1. Field of the Invention
This invention relates to a supersonic flight vehicle and more particularly to a supersonic engine bleed inlet for optimizing engine performance.
2. Description of Related Art
Aircraft engines which are designed to operate at supersonic speeds often have engine inlet configurations with bleed systems including bleed inlets that must operate in complex supersonic flow fields. Through viscosity, friction extracts momentum from the flow (boundary layer) nearest the wall. Portions of the boundary layer experience momentum reductions large enough to be incompatible with momentum requirements further downstream. Unless such deficient momentum air is removed or re-energized, the boundary layer subsequently separates. Such separation reduces the total pressure recovery, increases flow distortion levels, and can lead to unstable flow conditions. In order to compensate for this problem, engine inlets have to be sized larger than would otherwise be required for a non-separated engine inlet flow. Therefore, it is very desirable to have an engine inlet that can operate at high Mach number supersonic flow without boundary layer separation, in order to operate safely and efficiently and to be able to design and operate lower weight supersonic aircraft engines.
It is well known in the supersonic propulsion field, to use inlet boundary layer bleed systems, for removing boundary layer flow along the interior inlet walls for engines designed to operate at supersonic flight Mach numbers, in order to avoid these problems. While the use of such inlet boundary layer bleed systems results in improvements in both engine cycle performance and engine stability, a drag penalty is paid since such bleed air is ultimately discharged to ambient conditions at a lower exit momentum than its freestream level upstream of the bleed inlet. Optimizing thrust minus drag at a given fuel flow helps optimize system performance. Therefore, it is also very desirable to have an engine inlet that can operate at high Mach number supersonic flow without boundary layer separation and also reduces the drag penalty due to bleed flow momentum losses.
This can is accomplished in the present invention by reducing the amount of bleed flow required to achieve the increased total pressure recovery and reduced airflow distortion. Test data on typical boundary layer bleed systems indicate that in order to achieve the desired pressure recovery increase and distortion reduction, the amount of boundary layer which must be removed is more than twice the amount which was initially momentum deficient. Major contributing factors to the need for this excessive amount of bleed are the pressure disturbances originating downstream at the shockwave interaction, which propagate upstream through the subsonic portion of the boundary layer. Such pressure disturbances distort the upstream boundary layer flow profiles, which then become a source for additional losses and distortion in the main stream flow above the boundary layer.
There is a great and long felt need among supersonic aircraft and aircraft engine designers for a supersonic inlet bleed system that can operate efficiently to prevent boundary layer separation with a minimum amount of boundary layer bleed flow and drag. The present invention is directed towards such an inlet bleed system.